The present invention relates to passive shock load isolation apparatus suitable for use on spacecraft launch systems.
Spacecraft launch systems present severe requirements for shock load isolation apparatus. Shock loads are typically large magnitude disturbances having vibration energy at 100 Hz and above and can be of a transient or continuing nature, and can have both spacecraft axial and lateral directional components. During launch and during spacecraft separation from the launch vehicle, substantial shock load vibration energy is transmitted from the launch vehicle to the more delicate spacecraft. This loads environment drives up the size and strength requirements of the spacecraft and spacecraft-to-launch vehicle interface hardware. In the development of shock load isolation apparatus, as with any space-flight hardware, the mass and space requirements of the apparatus must be aggressively minimized so to reduce the cascading design impact to supporting hardware and thereby reduce the associated additional development and launch costs. Additionally, the overall system dynamics of the spacecraft and launch vehicle system must be predictable and carefully controlled in design so to allow controlled and stable flight. With the desire to launch a wide array of payloads comes the need for shock load isolation apparatus that can be easily scaled up or down to accept the broadly ranging size and mass of the spacecraft structures. And perhaps most importantly, the high costs and lengthy development times associated with satellites and other space bound hardware require that a premium be placed on obtaining durable and reliable spacecraft mounting hardware. Because the shock isolation mount, to be effective, must be the only connection of the spacecraft to the launch vehicle, it must meet these demands of high durability and reliability in concert with its shock isolation characteristics.
One described spacecraft axial vibration isolator is disclosed in U.S. Pat. No. 5,961,078. This patent claims that a payload attachment interface ring can be utilized to add axial compliance to the payload and launch vehicle interface by staggering the attachment points of the payload relative to the attachment points of the interface ring to the launch vehicle. This patent also suggests that damping can be added to the inherent damping of the payload attachment interface ring by making the ring a laminate which includes a `yieldable material to provide inherent hysteresis qualities`. The reality and deficiency of this concept for shock load isolation is that little additional axial compliance and deflection can be added without encountering stress failures in the payload attachment interface ring. The compliance gained is a simple addition of the beam bending and beam torsion compliances of the payload interface ring induced by staggering the payload attachment points relative to the base structure support pads. More importantly, the damping available within this concept from the bending and torsion of the interface ring material is commonly quite small, and provides relatively little damping to shock loads. Also, since the deflections of the interface ring are limited by the failure stress limits of the interface ring's load carrying material, the additional damping possible by making the ring a laminate including a yieldable material is severely limited. The more substantial axial compliance and damping characteristics necessary for effective shock load isolation may have been part of the `specially designed flexure feature` referred to in this patent's description, but it is neither described nor claimed.
A lateral force vibration isolator is claimed in U.S. Pat. No. 6,012,680. This patent discloses a concept of a payload attachment structure incorporating a circular array of beam flexures arranged to allow lateral compliance between the payload and supporting base. This concept provides no significant axial compliance nor damping necessary for isolation of the payload from axial shock loads.
U.S. Pat. No. 5,878,980 discloses a device where, to achieve sufficient damping, an elastomer is utilized in the primary load path within the mount of the spacecraft to the launch vehicle. This approach for obtaining damping has been commonly employed in the vibration isolation field, but it results in a mount with the distinct disadvantage of having a significantly lower strength and reduced fatigue life relative to the invention disclosed, which provides high damping but with a complete high strength material load path. The insertion of the elastomer into the primary load path in U.S. Pat. No. 5,878,980 also results in a mount with nonlinear load deflection behavior which complicates the design process and reduces the dynamic performance predictability of the launch system.